The present invention generally relates a method for fabricating and assembling a combustor liner in a turbine engine, and, more specifically, to a method for fabricating and assembling a multi-axial pivoting combustor liner in a gas turbine engine.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and burned for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. Combustors used in aircraft engines typically include a combustor liner to protect surrounding engine structure from the intense heat generated by the combustion process.
A conventional combustor liner has a cylindrical shape with one open end. A thin sheet metal material, capable of withstanding high temperature conditions, is usually used to fabricate the body through a forming process. The liner is often supported on one end or suspended by a few points. The conventional liner assembly and fabrication technique is adequate only for low cycle and low performance engines.
U.S. Pat. No. 3,613,360, referring to FIG. 1, discloses a radial combustion chamber having an outer case (14) enclosing an inner casing (15, 20) to form an annular combustion chamber. An air plenum or passage extends along the outer side of the combustion chamber. Multiple flanges are bolted together to form a combustion chamber. Bolts and threaded assemblies are used to make the joints, there is no pivoting features present in the combustion chamber.
U.S. Pat. No. 4,614,082 discloses a radial combustor liner having a plurality of panels mounted by means of a slideable friction mounting arrangement upon a high strength structural frame. Bolts are relied upon to fasten the aft end and the free support at the forward end. Tongue grooves are mainly used for engaging panels that make up the liner. Thus, the resulting combustor lacks any pivoting features during operation.
U.S. Pat. No. 6,434,821 discloses a method of fabricating an annular radial liner for a combustion chamber. The patent is concerned with improvements in the strength and durability of the liner by using a large forging with several joggles and scallop to ease thermal stress (12). See FIG. 1. Sheet metal members are jointed and welded into an annular section (50, 52 and 54). The aft end of the liner is bolted to the case in both the axial and radial directions (34). Therefore, the large radial liner has limited axial sliding with no multi-axial pivoting capabilities.
As can be seen, there is a need for an improved method for making a multi-axial pivoting combustor liner for gas turbine engines. A combustor liner made by such an improved method must have the ability to control small amounts of air leakage, provide easy assembly, have no flow path steps, and tolerate thermal and mechanical stresses while minimizing thermal wear and fretting for the life of the liner.